1. Field of Invention
The present invention relates generally to blade assemblies for rotary wing aircraft, and more particularly, but not by way of limitation, to an improved blade assembly for rotary wing aircraft wherein the blade may be controllably twisted to various positions in flight to optimize the angle of attack along the span of the blade for various flight conditions.
2. Description of Related Art
The lift developed by a conventional aircraft wing depends primarily on two factors: the angle of attack of the wing and the velocity of the air in relation to the wing. In the case of a rotary wing aircraft such as a helicopter, the air velocity in relation to the wing is produced by a combination of aircraft motion and the rotation of rotor blades.
The air velocity over a particular section of the blade, referred to herein as a blade element, is a function of the distance of the blade element from the rotational axis of the rotor. In other words, air velocity is greater at the outer end of the blade than it is near the rotational axis of the rotor since the speed of the outer end of the blade is greater than the speed of the blade near the rotational axis of the rotor. As a result, each blade of a rotor assembly may be formed so that it is twisted about its longitudinal axis in order to maintain a favorable angle of attack along the span of the blade.
Helicopter and tilt-rotor blades are supported in bearings that allow the blades to be rotated about their longitudinal axis. This rotation allows the pilot to make a gross change in blade pitch, either collectively or cyclically, in order to control the total lift and horizontal flight direction of the aircraft. This rotation of the entire blade does not, however, allow each blade to be variably twisted so as to set each blade at an optimum pitch at each blade element along the span of each blade for a specific flight condition.
The optimum performance of a rotor blade assembly occurs when the angle of attack of each blade element along the span of each blade is related to the angle of other blades in such a way as to provide a consistent inflow/outflow velocity across the rotational plane of the blades. This requires each blade element along the span of each blade of the rotary blade assembly to have the proper angle of attack in relation to the air velocity at that particular blade element. For a rigid blade, the twist distribution is constant regardless of the flight conditions. Thus, the twist can be optimized for only one flight condition. For example, if hover performance is most critical, blade twist is optimized for airflow conditions that prevail during hover, with the consequence of degraded performance in all other flight conditions. Similarly, optimization of blade twist for high speed cruise degrades performance in all other flight conditions, including hover.
In addition, the ride quality of a helicopter with a fixed blade pitch distribution is degraded by the non-optimum angle of attack of the individual blade elements from root to tip, and by the inability of relatively more rigid blade systems to respond to gusts or other turbulence through which the helicopter may be flying. If the blade angles can respond to these varying flight conditions in such a way as to minimize short-term fluctuations in spanwise lift, the ride quality will be improved throughout the range of the flight conditions.
As an alternative to a blade having a fixed or constant twist, a blade which can be variably twisted about the longitudinal axis of the blade can provide optimum blade pitch at each blade element along the span of the blade for different flight conditions. However, in order for the blades to have the potential for variable twist, they must be designed with a controlled amount of torsional flexibility. Unfortunately, torsional flexibility brings with it the disadvantage of making the blades more susceptible to torsional flutter, a dynamic interaction between aerodynamic, flexural and mass properties of the blades. Torsional flutter can be described as a self-sustaining torsional vibration of the blades in which the blade pitch varies above and below acceptable limits at relatively high frequency and in such a way as to decrease the aerodynamic performance of the blade and impose high structural loads on the blade.
To this end, a need has long existed for an improved rotor blade assembly which allows a torsionally flexible blade to be controlled in flight to optimize flight performance while also diminishing the likelihood of flutter. It is to such an improved rotor blade assembly that the present invention is directed.